Stabilizer for rotary wing aircraft



Dec. 18, 1951 F. A. ERICKSON 2,578,630

STABILIZER FOR ROTARY WING AIRCRAFT Filed Dec. 5, 1947 4 Sheets-Sheet l N 03 q A q I H R H Ln.

. INVENTOR. H FRANK A. ERICKSON ATTORNEY Dec. 18, 1951 F. A. ERICKSON 2 5 STABILIZER FOR ROTARY WING AIRCRAFT Filed Dec. 5, 1947 4 SheetsSheet 2 INVENTOR. FRANK A. ERICKSON ATTOQNEY Dec. 18, 195] F. A. ERICKSON STABILIZER FOR ROTARY WING AIRCRAFT 4 Sheets-Sheet 3 Filed Dec. 5, 1947 INVENTOR. ERICKSON FRANK A.

ATTORNEY Dec. 18, 1951 F. A. ERICKSON 2,573,539

STABILIZER FOR ROTARY WING AIRCRAFT INVENTOR. FRANK A. ERICKSON ATTORNEY Patented Dec. 18, 1951 UNITED STATES, rArgurqorrlcr-z 2,578,680 a V V STABILIZER FOR ROTARY WING AIRCRAFT Frank A. Erickson, Elizabeth City, N. G. Application December 5, 1947, Serial nmsaoos 13 Claims. (01. 24417.19) (Granted under the act of March 3,1883, as

. 1 a The invention described herein, if patented, may be manufactured and usedby or for the Government for governmental purposes without the payment to me of any royalty thereon in accordance with the provisions of the act of April 30, 1928 (Ch. 460, 45 Stat. L. 467).

My invention relates to rotary wing aircraft, and more particularly to stabilizers for such aircraft.

The rotary wing aircraft or helicopter is inherently unstable. in cruising flight. Attempts have been made to eliminate this inherent instability, mainly by the employment of vertical or, horizonta] stabilizers, or both, such stabilizers being mounted upon the fuselage of the aircraft, to directly control the attitude of the same. A disadvantage of this type of stabilizer lies in the fact that it must have a relatively large surface area in order to have an appreciable effect upon the attitude of the fuselage. Accordingly, the present invention is concerned with stabilizers which are relatively small in area and which will produce longitudinal, lateral and directional stability in helicopters through their effect on the cyclic pitch controls of the sustaining rotor, and rudder controls of the torque compensating rotor of the aircraft. The stabilizers embodying the invention are applicable to those helicopters in which longitudinal and lateral control is obtained by varying the pitch of the sustaining rotor blades during their cycle of rotation, and directional control by varying either the pitch of the torque compensating rotor, on single sustaining rotor types, or by diflferential control of the sustaining rotors in twin sustaining rotor types.

A primary object of this invention is to provide means for stabilizing rotary wing aircraft to maintain a given condition of flight through cruising speed range.

A further object of the invention is to provide a universal stabilizing system for helicopters which will produce suflicient stability in all reference planes so that the aircraft will maintain cruising flight conditions for several minutes, without using the cyclic pitch or rudder controls.

A further object is to provide a stabilizer unit to be attached to the swash plate of the cyclic pitch controls, and tending to reduce control sensitivity under cruising flight conditions to a degree comparable to that of fixed wing aircraft.

A further object is to provide stabilizers of the above mentioned type which provide sufficient stability to permit the use of a standard aeroplane type automatic pilot. I

A further object is to preload the cyclic pitch amended April 30, 1928; 370 0. G. 757) control system by the mounted upon the swash plate, thereby dampen-;

action ing cyclic pitch stick vibrations.

independently ofthe longitudinal and lateral stabilizer unit of theswash plate, to obtain directional stability in either single sustaining rotor types with a torque compensating rotor, or in twin and multi-rotor types employing differential action between the sustaining rotors to obtain directional control.

.A further object is to provide meatball stabilizer for use on the single sustaining rotor ype helicopter, having a torque compensating roton which can beset to automatically increase torque compensating rotor pitch, to compensate,

for torque in hovering flight, and also to automatically reduce torque compensating rotor pitch to that required for cruising flight.

A st lll further object of the invention is to provide a stabilizing system for rotary wing aircraft which will not interfere with precise hovering or slow speed maneuvers of the aircraft.

Other objects and advantages of the invention will become apparent during the course 0f the following description:

In the accompanying drawings, forming a part of this application, and in which like numerals are employed to designate like parts throughout the same,

. Figurel is a side elevation of a conventional single sustaining rotor helicopter, equipped with of the rear portion of the fuselage, part in h orizontal section,

Figure 4 is an enlarged fragmentary perspective view of a swash plate, and longitudinal and lateral stabilizers mounted thereon,

Figure 5. is a vertical longitudinal section taken on line 55 of Figure 4,

Figure 6 is a vertical transverse section taken" on line 6-6 of Figure 4,

view showing a screw drive for trim-tabs,

Figure 8 is an enlarged fragmentary side el-Q vation of a gear box and crank for. controlling the movement of trim-tabs, part in sectiom i of a stabilizerv Figure 7 is an enlarged'frag mentary sectional Figure 9 is a plan view of a modified form of longitudinal and lateral stabilizing unit,

Figure 10 is a vertical longitudinal section taken on line Iii-ID of Figure 9,

Figure 11 is a vertical transverse section taken on line of Figure 9,

Figure '12 is a side elevation of a modification of the invention showing'stabilizers applied'to a twin sustaining rotor helicopter, the fuselage being shown in vertical longitudinal section,

Figure 13 is an enlarged fragmentary plan view,

partly diagrammatic, illustrating the controls connecting the twin sustaining rotors, parts omitted, and,

Figure 14 is a fragmentary plan view'of elements of the pilot's lateral controls. I

My helicopter has a rotor provided with hinged blades, which rotor describes a cone whose extended axis of rotation passes through the center of gravity of the supported aircraft structure. The directional movement of the helicopter is obtained by tilting the axis of rotation of. the cone in the direction in which .the movement is desired. The tilting of the cone is accomplished by cyclical variation in the rotor blade pitch. The cyclical pitch changing mechanism includes an upper rotating swash plate section connected with the blades to progressively change their pitch during each cycle of rotation, and a lower swash plate which has a universal mounting and is angularly adjustable in any direction. This is conventional construction and is broadly shown in Patent 2,415,148. The present invention embodies a helicopter having a universally tiltable sustaining rotor with an upperrotatable swash plate section connected with the blades of the sustaining rotor to progressively-vary their pitch during each cycle of rotation, and a lower-universally tiltable non-rotatable swash plate section which is angularly adjustable to tilt the upper swash plate section. a

' Inthe drawings, where for the purpose of illustration are shown preferred embodiments of the invention, attention is called first to Figures 1 to 8 inclusive, wherein the numeral designates a conventional single sustaining rotor type of helicopter, having a fuselagel5, and including a sustaining or lifting rotor I6 and a torque compensating rotor I1 having variable pitch'blades H. The sustaining rotor lfiincludes three variable pitch rotor blades |8 which have their pitch changed during each cycle of rotation, through the action of a conventional cyclic pitch'control system. This cyclic pitch control system includes a swash plate assembly l9, having a lower nonrotatable swash plate section 20 and 'an upper rotatable swash plate section 2|. Generally vertical push-pull control rods 22 and 23 are connected to the lower swash plate section 20 in the usual manner and are adapted to be shifted longitudinally for tilting the swash plate section 20 longitudinally or laterally, with respect to the fuselage l5. V

Disposed beneath the lower swash'plate section 20, and rigidly secured thereto by means of a circular group of bolts 24, or the like, is a rigid stabilizer mounting plate 25, including outwardly laterally projecting portions or arms 26 and 21. The'arms 26 and 21 are bent downwardly slightly, adjacent to the inner portionof the plate25 beneath swash plate section 20,-and' upwardly at outer points 26" to form outer lateral inclined portions 28 and 29. A permanent or built-in dihedral angle is thus formed in the'stabilizer mounting'plate 25. The outer portions "28 and 29 of mounting plate 25 are elongated and tapered toward their outermost ends, and these outer portions 28 and 29 have rigidly mounted thereon generally horizontal rectangular airfoils or stabilizer vanes 30 and 3|. The outer portions 28 and 29 extend into and pass laterally through the airioils 30 and 3|, as shown, and are suitably rigidly anchored within the same. The portions 28 and 25 of mounting plate 25 are arranged near the frontal edges 32 and 33 of the airfoils, as shown in Figure 4. The left hand airfoil 33 is slightly wider than the right hand airfoil 3|, so that it may provide slightly greater lift to compensate 35 and 3| carry adjustable trim-tabs 34 and 35,

pivotally connected to the rear ends of airfoils 3G and 3|, and adapted to swing vertically with respect to the same.

Means are provided for manually adjusting the trim-tabs 34 and 35 from the cockpit. Such means include a gear box or housing 35, mounted in a convenient location for the use of the pilot,

preferably in the upper part of the cockpit, as

shown in Figure 1. Arranged within the housing 36 is a gear 31, mounted upon a transverse shaft 35, journaled in the sides of housing 36. A manually operated crank 39 is secured to one end of shaft 38, and is disposed on one outer side of the housing. A bevel gear 40 is mounted upon shaft 38, and is in meshed engagement with a bevel gear 4|, secured to the adjacent end of a flexible drive shaft 42. The bevel gears 40 and 4| are disposed outside of housing 36. A triangular frame or yoke 43 is arranged within housing 36, and pivotally connected to the sides of the housing through a pin 44, or the like. Rotatably mounted upon the frame 43 at one inner corner of the same is a gear 45, connected with a shaft 45 which rotates with gear 45. A bevel gear' l'l is mounted upon shaft 46 to rotate therewith. The bevel gear 41 meshes with a bevel gear 48 secured to the adjacent end of a flexible drive shaft 49. The bevel gears 41 and 48 are disposed outside of housing 36. Rotatably mounted upon the opposite inner corner of frame is an idler'ge'ar 50, which is in permanent meshing engagement with the gear 45, as shown. A'gear shifting handle 5| is provided, and rigidly connected to the frame 43. This handle serves as means for swinging the frame 43 aboutpin 44 for shifting gears 45 and 50 into and out of engagement with gear 31. The handle 5| operates in a slot 52 in housing 36, as shown.

The flexible'shafts 42 and 49 extend rearwardly and upwardly toward the airfoils 30 and 3|, Figure l, and are arranged near the inboard edges of the airfoils. Suitable brackets 53 and 54 are rigidly secured to the inboard edges of airfoils '35 and 3|, near their frontal edges 32 and 33, as shown. Rigidly held in place by each of the brackets 53 and 54 is a longitudinally extending internally screw threaded sleeve 55, in which is rotatably mounted an externally screw threaded rod- 55. The rod 55 is turnable, and will travel longitudinally within sleeve 55 in either direction. The rod 56 has a forward socket 57 into which the adjacent end of one flexible shaft is secured for rotation with the rod'56. Each rod 55 has a rearwardly longitudinally extending portion 58 connected through a ball and socket joint 59 with a longitudinally extending connecting rod 60, in turn pivotally connected at 5| to a depending armor tab 62 angelic together constitute a, single compact stabilizing unit for the lower swash plate section 2|]. The airfoils 30 and 3| have arelatively small total surface area, because they need only be capable of producing a change in the attitude of the swash plate sectionifl, and not inthe entire fuselage of the helicopter. ,In some installations, the total combined areas of both airfoils 30 and 3| are approximately only 5% square feet.

Arranged at the rear end of the fuselage l5 and laterally opposite w the torque compensating rotor ll isa vertical directional stabilizer vane or airfoil 63. This airfoil 63 has a frontal edge 64, and the airfoil is pivotally mountedupon a generally vertical shaft 65 and adapted to swing horizontally upon the same. The shaft 65, carrying the airfoil 63 is rigidly connected to the rear end of the fuselage, andto the torque compensating rotor gear box or housing 66, by means of suitable rigid brackets 61 and 68 respectively. Rigidly secured to the inboard side of airfoil 63, and arranged generally at right angles to the same and extending laterally thereof is an arm 69, pivotally connected atll] with a push-pull rod N. This rod 11 extends into the fuselage and projects forwardly longitudinally in the same and is pivotally connected at 12 with a transverse crank-lever 13. A retractile coil spring 14 is connected at its forward end to the fuselage structure, and at itsrear end 15 it is connected to the end of crank-lever 13 adjacent to rod H. The transverse crank-lever 13 has a fixed pivot 16, near its longitudinal center, and is adapted to be swung horizontally about such pivot. At its end remote-from rod H, cranklever i3 is connected to a cable 11, which is tied into one run 18 of the manually operatedpitch control cable 11 for the torque compensating rotor 17. Cable TI is tied into the run 18' at 18 as shown. Laterally inwardly of spring 14, crank-lever '13 is connectedto a cable 19, which is tied into the opposite runl9 of cable 11, as at 80. The cable ll" is wound upon a spool 8|, mounted upon a rotatable transverse horizontal shaft 82. The cable H is so wound upon spool 8| that when one of its runs is wound up on the spool, the otherrun is let out therefrom. The runs 18 and 19' of cable'll extend forwardly in the fuselage of the helicopter and are connected in a conventional manner to conventional manually operated foot pedals in the cockpit, which are actuated to vary the pitch of the blades ll of torque compensating rotor ll. The rotatable shaft 82 carries a sprocket wheel 83 at its outer end, and a sprocket chain 84 engages about the sprocket wheel 83, and is also operatively connected with a sprocket wheel 85 mounted upon a rotatable shaft extension 86, which changes the pitch of blades I1, when rotated. In operation, the horizontal airfoils or stabilizers 3B and 31 provide both lateral and longitudinal stability through their reaction upoh; the lower swash plate section during forward flight. The airfoils 3t! .and 3| placeaload in the cyclic pitch control system which has the effect of dampening out stick vibrations during cruising flight. The trim-tabs 34 and 35 are actuated from the cockpit'by means of "the crank 39 and gearing shift handle 5|. Longitudinal trim is obtained bymoving the trim-tabs together, and lateral trim bymovlng them different a lr- This is a com lish d by turning gee cockpit.

39, and swinging gear shifting handle 5i, In

order to swing trim-tabs. and 35 togethen and.

in the same direction, handle 5| should be swung to the left, Figure 8, so that idler gear 50 meshes with gear 31. To move the trim-tabs together in opposite directions, the handle 5| shouldbe swung to its right position shown in Figure 8. The trim-tabs may be moved into any relative;

i position found to be necessary. The airfoils 30 and 3| have practically no effect during hovering flight, and do not interfere with controllability of the helicopter at that time. All forces causing unbalance in the cyclic pitch controls tend to be balanced out by the airfoils 30 and'3l." The built-in dihedral angle of these airfoilsproduces' added lateral stability. It is believed that'the' addition of the airfoils 30 and 31 is enough to provide sufficient stability and dampening effect to permit the use of a standard aeroplane-type automatic pilot. Directional stability and torque compensation are obtained by means of the vertical airfoil 53, which is acted upon by spring M, tending to apply. full torque in hovering flight. The airfoil 63 could be arranged anywhere on the fuselage where the airflow is not disturbed, but the preferred location is shown and described. Spring 14 places a con-' stant load on the air foil 53, tending to applyfull left rudder control by increasing the pitch of the torque compensating rotor blades H. The inflow of air past airfoil to the torque compensating rotor I! aids the spring 14 in holding left rud-' der control in hovering flight. In forward flight, airfoil 63 tends to align itself with the relative wind and opposes the action of the spring 14. The effectiveness of this action of airfoil, 63 increases with increase in speed. At intermediate speeds spring i l will apply a small degree of left rudder control to compensate for the reduced torque reaction. At high speeds it will practically feather the torque compensating rotor, as little or no torque compensation is required at high speeds. When flying speed is reduced, spring M partially overcomes the effect of the airfoil, as the airfoils effectiveness is reduced with the reduced flow of air over it. In hovering flight, the airfoil will cut out completely, allowing the spring to apply full left rudder control.

The helicopter fuselage is affected by side wind gusts in the same manner as a wind sock, that is, it will swing with the nose of the aircraft into the gust. The inertia of airfoil 53 being much less than that of the fuselage permits the airfoil to be deflected by the gust before the fuselage is deflected. Assuming that a wind gust strikes the right side of the aircraft in cruising flight, airfoil 63 will be deflected to the left, looking forward in Figure 3, rod ll will be shifted longitudinally forwardly turning crank-lever l3 counterclockwise. Crank-lever 13 in turn will move the directional controls through cables 11 and TI, increasing the pitch of torque compensating rotor blades [1, which opposes the side wind gust, tending to prevent the aircraft from turning into the Crank-lever "53 moves cables 11 and 11', which causes drum 8| and shaft 82 to rotate, driving sprocket chain 34, and shaft extension 86, to change the pitch of torque compensating rotor blades H. The separate runs 18 and 19' of cable 'l'l may always be manually operated by the foot pedals in the In addition airfoil 63 automatically moves the runs 18 and 19 through the medium of crank-lever 13, and associated elements. .Attention is called next to Figures!) to 11 in;

estates 7. elusive, mean I have shown a modified for-m df lateial and longitudiilal stabilizing unit, Similar to the lateral and longitudinal Stabilizing unit "first form of the invention, which includes airfoils 30-and 31. In this form of the invention, I contemplate omitting trim-tabs "3'4 and 35. In Figures 9 to 11 inclusive, generally horizontal rectangular stabilizers or airfoils '30 and 3i are provided, and these stabilizers are carried by a rigid'niounting plate 25 similar to the-mounting plate 25. 'The mounting plate '25 is bolted to the bottom of a lower swash plate section in the sameman'ner as the p1ate' is bolted to the swash plate section 20. "in this form of the invention, the trim ta'b's are omitted, as stated, and the airfcil's 3'0 and'3'1' are adapted to swing bodily verticallygthrough a limited travel. The mounting plate 25 "projects laterally upon both sides of swash plate section 20' and terminates adjacent to the inboard edges of stabilizers and BI, assl'iown. Cylindrical bars or shafts 8S and 90 are provided and are permanently rigidly'connected at their inboard ends to the outer er'i'ds of mounting plate 25. The shafts t9 and 00 are non-rotatable with respect to mounting plate 25', and project into and through the stabilizers 30 and 31', which are rotatably mounted upon the shafts 8'9 and '90, and adapted to be swung vertically bodily thereon.

' Flexible shafts 42' and "49 are provided and are connected with the gear box 36 in the same manner "shown and described in connection with the first form of the invention. Each of the airfoils '30 and 3| has a depending bracket SI rigidly mounted upon its associated relatively stationaryshaft 09 or 90, and including an inclined right angle extension 92, which projects beneath the lower side of the airfoil and inwardly. Rigidly secured to each extension "92 is an internally screw threaded sleeve 03, receiving an externally screw threaded rod a l, similar to the rod 58. Each rod "94 is connected through a ball and socket joint 95 with a. connecting rod-9G, in turn -pivotally connected with a depending lug 91, rigidly attached to the associated stabilizer 30' or 3|.

In the form of the invention shown in Figures 9 to 11, when the crank '39 of gear box 30 is turned, the flexible drive shafts 02"and49 are turned, causing the screw threaded rods '94 to described in the first form of the invention, the

stabilizers 30 and 'SIrnay be moved together or differentially. All other parts are identical with those shown and described in connection with the first form of the invention.

Attention is called next to Figures '12 and 13, wherein a modified form of the invention is i1"- lustr'ated, and the stabilizers are shown-applied to a. twin sustaining rotor ty'pe helicopter. 'In-Figures 12 and 13, the numeral 98 designates the fuselage of a twin sustaining rotor helicopter, having forward and aft sustaining rotors 9'9 and I00, each of which is identical in construction'and operation to the rotor I6 shown and described in connection with'the first form of the invention. The rotors 99 and I00 are shown equipped with horizontal stabilizer units NH and I02, and "each of the units IOI and I02 includes airfoils I03 and I04 which are identical in construction and operation withthe airfoils 30 and .H. These airfo1ls'l03 and I04 are equipped with trim tabsand operating means therefor, identical to those shown and described in connection with the first form'of the invention. The stabilizer units [DI and I02 act u on the rotors 99 and I00, to afford longitudinal and lateral stability for such rotors in the same manner as the single rotor is stabilized.

I 'A vertical directional stabilizer I05 is disposed adjacent to the rear end of fuselage 98, and is mounted upon a vertical rock shaft I06, and adaptedto swing horizontally thereon. The vertical stabilizer I05 includes laterally projecting arms I 01 and I08 rigidly connected therewith, and the outer ends 'of'thes'e arms are pivotally connected to'cables I09 and H0. The cables I09 and H0 are in turn pivotally connected with the aft corners of a horizontally 'swingable sector III. A rod 112 is pivotally connected at one end to one side of sector III, as shown at -I I3, and this rod projects rearwardly from the sector, and is pivotally connected to the vertical arm N4 of a bell crank, whose generally horizontal arm I I5 is pivotally connected to a vertical push-pull rod H6 (corresponding to rod 23, Figure 4), in turn connected at its upper'end to one side of the lower swash plate section IIB' of rotor I00. The 'forward corners of sector III are connected to cables II! and I I8, which are parallel and extend longitudinally within the upper part of the 'fuselage and engage beneath pulleys H9, as shown. The forw-ardends of cables II! and H8 are connected to the aft corners of a horizontally swingable sector I20, having a push-pull rod I2'l pivotally connected to its side remote from the rod H2, and projecting forwardly from the sector, and pivotally connected to the vertical arm I22 of a bell crank having agen'er-ally horizontal arm I23. The arm I23 is'in turn pivotally connected with a generally vertical push-pull rod I 24 (corresponding to rod 23, Figure 4), connected at its upper ends'to one side of the swash plate I 24 of rotor 99. 'The push-pull rods I20 and H6 are connected to opposite sides of the swash plates of rotors 99 and I00. The swash plate sections H6 and I24 are swung longitudinally of the fuselage by rods 22, corresponding to red 22, Figure 1-, and rod's2 2' and 22 are'mo'vedin the conventional manner.

Manually operated cyclic pitch controls are .provided, and these controls are adapted to change the attitude of rotors 99 and I00 in unison to obtain lateral and longitudinal control for the helicopter. Inasmuch as both lateral and directional controls are obtained by tilting the swash plates laterally (in unison'for lateral control and diiferentially for directional control), the complete lateral and directional control systems are shown schematically. The lateral control consists of a. stick I 2 5, universally pivoted at I26, and having a depending extension I 21, ivotally connected to a transverse connecting rod I27, in turn pivotally connected to a pivoted bell crank I28 which-swings horizontally. The bell crank I28 is pivotally connected with a longitudinally shiftabl'e rod I29. The rod I20 is pivotally connected to the top 'end of a walking beam I30, which is pivoted near its center, as at -I3|, and pivotally connected at its lower end I32 to a longitudinally 'shiftable rod I33. The rod I33 extends longitudinally rearwardly and is disposed near the bottom of the fuselage, and is pivotally connected at its rear end to the vertical arm I34 01 abell crank, whose generally'horizontal arm I35 is pivotally -connected to a vertical push-pull rod I36, in turn pivotally "connected at its upper end to the horizontal arm I31 of a bell crank including a generally vertical arm I38, connected to a longitudinally shiftable long push-pull rod I39.

The rod I39 serves to connect the rotors 93 and I with the manual controls including stick I25,

rod I44 adapted to be shifted vertically within a .fixed guide I45, and pivotally connected at its top end to the fulcrum of the bell crank including arm I22. The rod I44 is adapted to shift the bell crank including arm I22 bodily vertically.

Rudder controls are provided, including foot pedals I 46; which are pivoted at I41, and connected at their lower ends I48 to longitudinally shiftable cables I49. The cables I49 extend rearwardly, and pass about fixed pulleys I49. These cables extend upwardly, forming generally vertical runs I50, which pass about fixed pulleys I50. The runs I50 are attached to the forward corners of sector III, as shown.

The operation of the form of the invention shown in Figures 12 and His as follows Assuming that a side gust of wind were to strike the rear of fuselage 98, tending to deflect it to the left, the vertical stabilizer I would be almost immediately. deflected to the left, movin cables I 09 and I I0, which in turn swing the quadrant III. The motion imparted to the quadrant I I I is transmitted through push-pull rod I I 2, bell crank including arms H4 and I I5, and push-pull rod I I6 which tilts the swash plate section I I 9 of rotor I00 laterally to the right. Likewise, simultaneously, the swash plate section I24 of forward rotor 99 is tilted laterally to the left, through the action of cables Ill and H8, which transmit the motion in quadrant I I I to quadrant I20, which in turn moves push-pull rod I2I, bell crank I22 and push-pull control rod I24. Thus when the wind gust displaces the stabilizer I05, such stabilizer affects an immediate change in the attitude of both sustaining rotors 99 and I00, utilizing their differential cyclic pitch control system. Since the fuselage 98 has far greater inertia than stabilizer I05, the stabilizer will move, and produce the stabilizing correction of the rotors 99 and I00, before the fuselage can be moved appreciably by the wind gust.

The manual cyclic pitch controls including stick I25 are also connected in the system at all times. The stick I25 is swung laterallyabout its universal pivot I29 to shift rod I29 longitudinally, swinging Walking beam I39 about its pivot, shifting rod I33 longitudinally, turning the bell crank including arms I34 and I35, shifting rod I39 longitudinally. The rod I39 causes bell cranks I40 and I43 to turn simultaneously. The bell crank I40 shifts rod I4I vertically in fixed guide I42, causing bell crank including arm II4 to be shifted vertically bodily, shifting rod IIB vertically and tilting the swash plate section IIB of rotor I09 laterally. The bell crank I43 simultaneously shifts rod I44 vertically in the opposite direction to the movement of rod I4I, causing bell crank including arm I22 to shift vertically bodily with it, shifting control rod I24 swash platesection N6 of rotor I00. The directional controls including foot pedals I46 are operatively connected with the stabilizer I 05. Foot pedals I46 may be pivoted at I41, for shifting cables I49 longitudinally, shifting runs I50, which are attached to the forward corners of sector III, as stated. The sector III may thus be turned by manipulatingv the pedals I46, and the stabilizer I05 is thus manually movable.

The longitudinal and lateral stabilizer units I0! and I02 function to stabilize the swash plates of rotors 99 andI in the, same manner as the stabilizers 30 and 3| stabilize swash plate I9.

The basic idea embodying the stabilizer I05 is applicable also to co-axial sustaining rotor helicopters, in which the total pitch is changed differentially to obtain directional control.

It is to be understood that the forms of the invention herewith shown and described are to be taken as preferred examples of the same, and that various changes in the shape,-size and arrangement of parts may be resorted to, without departing from the spirit of the invention or the scope of the sub-joined claims.

Having thus described my invention, I claim:

1. In a helicopter, in combination, cyclic pitch controls including a non-rotatable swash plate section, a mounting plate secured to said nonrotatable swash plate section, .shafts fixedly secured to the mounting plate, stabilizer vanes pivotally mounted upon the shafts to vary their angle of attack, and means to change the angle of attack of said stabilizer vanes. j

2. In a helicopter, a fuselage, sustaining rotors mounted upon the fuselage, cyclic pitch control means for each sustaining rotor and including a non-rotatable swash plate section, and stabilizer vanes mounted upon each of said non-rotating swash plate sections.

3. In a helicopter, the combination with cyclic pitch controls including a non-rotatable swash plate section, of stabilizer vanes mounted upon the. non-rotatable swash plate. section and arrangedlat a, dihedral angle.

4. In a helicopter, a fuselage, sustaining rotors mounted upon the fuselage, cyclic pitch control means for each sustaining rotor and including a non-rotatable swash plate section, stabilizer vanes mounted upon each of said non-rotatable swash plate sections, and operating connecting means between the swash plate sections to cause them to tilt laterally in the same and opposite directions.

4. In a helicopter, a fuselage, sustaining rotors mounted upon the fuselage, cyclic pitch control means for each rotor including a non-rotatable swash plate section, stabilizer vanes mounted upon each of said non-rotatable swash plate sections, manually operated means connecting the swash plate sections and causing them to tilt in opposite directions, a directional stabilizer vane mounted upon the fuselage, and means operated by the directional stabilizer vane and connected with the manually operated means to move said manually operated means.

6. A helicopter, comprising a fuselage, a sustaining rotor mounted upon the fuselage, cyclic pitch controls for the sustaining rotor and including a non-rotatable swash plate section, stabilizer vanes mounted upon opposite sides of said swash plate section, and means to connect the stabilizer vanes with the swash plate section for control thereof.

If g

" 7. A' helicopter, comprising a'iuselage', a; sustaining rotor mounted upon the fuselage, cyclic pitch controls for the sustaining rotor and including a non-rotatable swash plate section,

stabilizer vanes arranged upon opposite sides of said swash plate section, means to connect the stabilizer vanes with said swash plate section,

a directional stabilizing vane mounted upon the fuselage, manually operated means connected with said swash plate section to tilt the same,

and connecting means between the stabilizer vane and" manually operated means for actuation thereof.

8'. A helicopter, comprising a fuselage, a sustaining rotor mounted upon the fuselage, cyclic pitch controls for the sustaining rotor and inrotatable' stabilizing vanes arranged upon opposite sides of the said swash plate section and connected with said swash plate section for controlling its action.

' means.

10. In combination, a helicopter provided with rotor blade pitch controls including a non-rotating swash plate section, stabilizing means including non-rotating airfoils mounted for movement about an axis offset from the center of pressure of said airfoil in response to aerodynamic forces, and means to connect the stabilizing means with said non-rotating swash plate section for actuation thereof.

11. In combination, an aircraft. having rotative sustaining means, control means connected to said sustaining means, non-rotating laterally eluding a lower non-rotatable swash plate section mounted for universal movement and nondisposed airfo'ils. mounted on said aircraftsaid mounting including means to permit said: airfoils to rock about an axis parallel tothe longitudinal axis of said aircraft. and to rock about an axis parallel to the pitching axis of said aircraft, means connecting said airfoils to said control means, the center of pressure of said airfoils being offset from the aforementioned axes about which said airfoils rock.

12. In combination, a helicopter provided with a sustaining rotor and cyclic pitch controls therefor and including a swash plate having a nonrotating portion universally mounted, stabilizing vanes laterally disposed from said swash plate, the spanwise axes of said vanes intersecting at a point below said swash plate, said axes being angularly related to provide a positive dihedral angle between the planes of said vanes, means connecting said varies with the non-rotating portion of the swash plate for operation of said swash plate in response to aerodynamic forces acting on the stabilizer vanes.

13. The combination asset forth in claim 12 and including trim tabs mounted on the stabilizer vanes and means for adjusting the angular relationship of the trim tabs with respect to the chord lines of said vanes.

FRANK A. ERICKSON.

REFERENCES CITED The following references are of record in the file of this patent:

UNITED STATES PATENTS Number Name Date 1,800,470 Oemichen Apr. 14, 1931 2,318,259 Sikoorsky May 4, 1943 2,368,698 Young Feb. 6, 1945 2,402,294 Pitcairn Jan. 18, 1946 2,443,192 Moeller June 15, 1948 FOREIGN PATENTS Number Country Date 155,974 Switzerland Oct. 1, 1932 844,048 France Apr. 11, 1939 

